The present invention relates to propulsion systems and, more particularly, to a linear gridless ion thruster, which combines an ionization stage from a gridded ion thruster and an acceleration stage from a closed-drift Hall thruster to take advantage of the strength of both thrusters without suffering from the weakness of either.
The Rocket Equation (Equation 1):                               Mf          Mo                =                  Exp          ⁡                      (                          -                                                Δ                  ⁢                                      xe2x80x83                                    ⁢                  V                                gIsp                                      )                                              Equation        ⁢                  xe2x80x83                ⁢        1            
shows that the ratio of payload or final mass (Mf) over initial mass (Mo) depends on the velocity increment (xcex94V) needed for a spacecraft, and the speed at which exhaust propellant leaves the propulsion system; also known as specific impulse (Isp), which is proportional to propellant exhaust velocity through the gravitational constant (g). That is, the amount of propellant needed to achieve this xcex94V is reduced if the Isp of the propulsion system is increased. For example, cryogenic chemical rocket motors such as the Space Shuttle Main Engine are capable of producing specific impulses of about 450 seconds. Chemical rockets employed for long-duration space voyages must use non-cryogenic propellants that yield lower performance ( less than 330 seconds).
Studies have shown that ideally, an engine that would be used as the primary source of propulsion for orbit transfer missions or for satellite station-keeping should produce an Isp between 1000 and 2000 seconds. Spacecraft propulsion systems for interplanetary missions may need to generate even higher exhaust velocities. To achieve the desired performance, a propulsion system must accelerate a propellant gas without relying on energy addition through chemical reactions.
One approach is the application of electrical energy to a gas stream in the form of electrical heating and/or electric and magnetic body forces. This type of propulsion is commonly known as electric propulsion (EP). EP can be categorized into three groups. Electrothermal Propulsion Systems electrically heat a gas, either with resistive elements or through the use of an electric arc, which is subsequently expanded through a nozzle to produce thrust. Electromagnetic Propulsion Systems use electromagnetic body forces to accelerate a highly ionized plasma. Electrostatic Propulsion Systems use electrostatic forces to accelerate ions. In addition to possessing suitable exhaust velocities, an EP system must be able to convert onboard spacecraft power to the directed kinetic power of the exhaust stream efficiently.
To show the benefit of EP systems over chemical systems reference is made to FIG. 1. FIG. 1 is a plot of the Rocket Equation showing the final-to-initial mass ratio for a number of missions that use conventional propulsion systems. Clearly the smaller the mass ratio, the more expensive a mission becomes. While missions to Low Earth Orbit (LEO), the moon, and Mars require significantly more propellant mass than payload mass when using chemical propulsion systems, this is not the case for EP systems due to their high Isp. This fact translates into significant cost savings for commercial, military, and scientific space missions.
FIG. 2 shows payload mass and fraction delivered to Geosynchronous Earth Orbit (GEO) as a function of trip time for EP and chemical propulsion systems assuming a moderate launch vehicle (Atlas IIAS) is used. FIG. 2 compares the performance given by a bi-propellant chemical rocket (Isp=328 sec), an arcjet using hydrazine decomposition propellant (Isp=600 sec), and a Hall thruster using xenon propellant (Isp=1600 sec). As FIG. 2 clearly shows, the amount of payload delivered to GEO increases with Isp and with trip time. The former is because the launch vehicle places a fixed spacecraft mass in LEO and as Isp increases, the amount of propellant needed for the transfer reduces. The mass that was used for propellant in the all-chemical spacecraft can now be used for payload.
A 15% increase in payload mass can be realized by simply using EP for North-South stationkeeping (NSSK) and using chemical propulsion for the LEO-to-GEO transfer. While the LEO-to-GEO trip takes longer with more of the transfer being done with EP, less propellant is required. Hence, the high-Isp EP system is used more for longer transfers, and more payload can be delivered to GEO.
This principle is being considered for the human exploration of Mars. NASA has now expressed an interest in developing the capability to send a crew to Mars within the next two decades. However, mission cost is a clear driver. Since the LEO-to-MTO (Mars Transfer Orbit) xcex94V is a significant fraction of the total mission xcex94V, and hence accounts for much of LEO initial vehicle mass, NASA has baselined the use of a Solar Electric Propulsion (SEP) stage to raise a chemically-powered Mars Transfer (MT) stage to a highly elliptic orbit around the Earth. Once the MT stage is in the proper orbit, the crew uses a small, chemically-propelled vehicle to rendezvous with it. Once the crew is in place and the MT stage has been certified to be fully operational, it separates from the SEP stage and ignites its engines for the trip to Mars.
EP""s resurgence in recent years is due both to the public""s interest in space exploration and money that be saved by commercial spacecraft developers. As illustrated above, the latter comes by virtue of the fact that EP""s large specific impulse means that it can accomplish a mission with less propellant than conventional propulsion systems. The recent successes of the Deep Space-1 and Mars Pathfinder missions have helped to renew the public""s excitement about space exploration.
The Mars mission scenario described above reduces both trip time (for the crew) and initial spacecraft mass by utilizing a high-performance SEP stage. The key to developing the SEP stage is the utilization of an engine that posses high specific impulse, high thrust efficiency, and a large range of specific impulse over which it can operate while maintaining high efficiency.
At first glance, a gridded ion engine appears to be ideal for the above application. Ion thrusters have very high specific impulses and efficiencies, and have a moderately large range of specific impulses over which they can operate at better than 50% efficiency. However, since such an engine will need to process hundreds of thousands or millions of watts of power, conventional gridded-ion thrusters are inappropriate given the size requirement such an engine would have due to its space-charge and grid erosion limitations.
On the other hand, conventional single-stage Hall thrusters possess high engine efficiency at moderately-high specific impulses. However, the ability to operate single-stage Hall thrusters with long life at very high specific impulses has never been demonstrated nor can ionization processes be decoupled from acceleration processes. The latter results in the strong interdependence of discharge current, discharge voltage, and propellant flow rate that limits the operational flexibility of these engines.
Furthermore, since ions are created at various spots along the ionization/acceleration region, not all ions benefit from the full accelerating potential of the discharge, resulting in a loss of engine efficiency. Moreover, the effect on engine life of placing 1000-2000 V discharge voltages on single stage Hall thrusters (e.g., on the anode from back-streaming electrons) is unknown. Lastly, for specific impulses of xcx9c1300 seconds or less, conventional Hall thruster efficiencies are low because of the coupled ionization and acceleration zones. This would serve to limit the xe2x80x9cthrottlingxe2x80x9d capability of the SEP stage (e.g., to provide xe2x80x9chighxe2x80x9d thrust at moderate specific impulse for certain phases of its orbital burn).
The desire for high throttling performance (also known as xe2x80x9cDual Mode Operationxe2x80x9d) applies to a number of commercial, military, and scientific missions. For commercial and military satellites, for example, the high-thrust, lower-Isp mode would be used for LEO-to-GEO transfer while the lower-thrust, high-Isp mode would be used for station-keeping.
In single-stage Hall thrusters, shown schematically in FIG. 3, ions are accelerated by the electric field established between a downstream cathode and an upstream anode. An applied radial magnetic field in an annular discharge chamber impedes the motion of migrating electrons. The crossed electric and magnetic fields create an azimuthal closed electron drift; the Hall current.
Propellant is injected at the anode and collisions in the closed drift region create ions. The ionization and acceleration processes in such a configuration are closely linked, limiting the useful operating range of the thruster to around 2500 s specific impulse and  less than xcx9c60% efficiency. Operation below these values results in intolerable decay in thruster efficiencies ( less than 35% efficiency around 1200 s specific impulse). This prevents Dual Mode Operation from becoming a reality.
Ionization and acceleration can be made more independent by the introduction of an intermediate electrode in the channel; a two-stage Hall thruster. FIG. 4 is a schematic of a traditional two-stage Hall thruster. The intermediate electrode acts as the cathode for the ionization stage and the anode for the acceleration stage. This allows the ionization stage to operate at high currents and low voltages resulting in higher propellant utilization (the efficiency at which propellant atoms are converted to thrust-producing beam ions) and the acceleration stage to operate at variable voltages resulting in a wide specific impulse range of operation.
Overall thruster efficiency is enhanced in this configuration, as Equation 2 illustrates:                               η          t                =                  1                      1            +                                          I                d                            ⁢                                                V                  d                                /                                  I                  a                                            ⁢                              V                a                                                                        Equation        ⁢                  xe2x80x83                ⁢        2            
where xcex7t is the overall efficiency, I is current, V is voltage, and the subscripts xe2x80x98axe2x80x99 and xe2x80x98dxe2x80x99 refer to the acceleration and discharge (ionization) stages, respectively. Thus, efficiency is increased for low discharge voltages and high acceleration voltages.
Work by Tverdokhlebov on a two-stage anode layer thruster demonstrated high efficiency ( greater than 67%) at high acceleration voltages ( greater than 500 V), but was unable to lower the discharge voltage below 50 V because backstreaming electrons were not of sufficient energy to maintain the discharge. Therefore, in such a configuration the ionization and acceleration processes are still weakly coupled due to the dependence of the discharge on backstreaming electrons. Further, operation of such a thruster has not yet been shown to be efficient at powers less than 6 kW and below 2500 s specific impulse. It is clear that a configuration that does not depend on backstreaming electrons is warranted so that discharge voltages may be minimized and ion production costs lowered.
Researchers in Japan have shown that use of an emitting intermediate electrode significantly increased the efficiency of a two-stage Hall thruster. These results are shown in FIGS. 5a and 5b. The two-stage device with cathode heating outperformed its single- and double-stage (no cathode heating) operation. Note that the efficiency of this device is very low, but this is believed to be caused by poor design owing to a long channel length and not to any physical constraints.
The trends demonstrated in FIG. 5 may indicate that an emitting intermediate electrode will increase overall efficiency. However, neither the Japanese work referenced here or any previous work known to the inventors have used magnetic fields expressly designed for the purpose of enhancing ionization; a technique commonly used in ion engines with great success.
As the following discussion will show, a gridless ion thruster that utilizes the ionization efficiency of a gridded ion thruster with the acceleration processes of a Hall thruster appears to be ideal for the application described above.
The present invention is directed towards a linear gridless ion thruster (LGIT) for use as an ion source that can be used for spacecraft propulsion or plasma processing. The LGIT is composed of two stages: (1) an ionization stage composed of a hollow cathode, anode, and cusp magnetic field circuit to ionize the propellant gas; and (2) an acceleration stage composed of a downstream cathode, upstream anode, and a radial magnetic field circuit to accelerate ions created in the ionization stage. The LGIT replaces grids used in conventional ion thrusters (Kaufman guns) to accelerate ions with Hall-current electrons as is the case with conventional Hall thrusters.
Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples, while indicating the preferred embodiment of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention.